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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 200,000
Max Cl/Cd: 66.85 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe113-il-200000-n5.txt
Download as CSV file: xf-goe113-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3298   0.08721   0.08400  -0.0167   1.0000   0.0205
  -8.000  -0.3287   0.08384   0.08065  -0.0179   1.0000   0.0215
  -7.750  -0.3283   0.08050   0.07735  -0.0192   1.0000   0.0220
  -7.500  -0.3285   0.07716   0.07404  -0.0204   1.0000   0.0222
  -7.250  -0.3300   0.07384   0.07077  -0.0215   1.0000   0.0223
  -7.000  -0.3334   0.07068   0.06766  -0.0223   1.0000   0.0224
  -6.750  -0.3356   0.06731   0.06434  -0.0237   1.0000   0.0224
  -6.500  -0.3346   0.06357   0.06062  -0.0258   1.0000   0.0225
  -6.250  -0.3317   0.05970   0.05677  -0.0279   1.0000   0.0225
  -6.000  -0.3271   0.05570   0.05276  -0.0298   1.0000   0.0226
  -5.750  -0.3210   0.05170   0.04874  -0.0314   1.0000   0.0226
  -5.500  -0.3137   0.04767   0.04466  -0.0326   1.0000   0.0226
  -5.000  -0.3061   0.05056   0.04695  -0.0402   1.0000   0.0116
  -4.750  -0.2890   0.04627   0.04250  -0.0414   1.0000   0.0113
  -4.500  -0.2693   0.04216   0.03821  -0.0424   1.0000   0.0111
  -4.250  -0.2397   0.03756   0.03334  -0.0448   0.9980   0.0111
  -4.000  -0.1982   0.03283   0.02818  -0.0479   0.9940   0.0124
  -3.750  -0.1634   0.02814   0.02306  -0.0512   0.9894   0.0132
  -3.500  -0.1270   0.02474   0.01925  -0.0534   0.9845   0.0137
  -3.250  -0.0899   0.02213   0.01624  -0.0556   0.9796   0.0147
  -3.000  -0.0544   0.02022   0.01396  -0.0571   0.9728   0.0176
  -2.750  -0.0181   0.01781   0.01101  -0.0581   0.9658   0.0207
  -2.500   0.0165   0.01624   0.00902  -0.0590   0.9570   0.0245
  -2.250   0.0499   0.01519   0.00784  -0.0603   0.9463   0.0289
  -2.000   0.0836   0.01436   0.00673  -0.0612   0.9337   0.0342
  -1.750   0.1166   0.01338   0.00569  -0.0623   0.9187   0.0372
  -1.500   0.1503   0.01285   0.00508  -0.0634   0.9008   0.0411
  -1.250   0.1824   0.01221   0.00435  -0.0641   0.8785   0.0410
  -1.000   0.2134   0.01171   0.00372  -0.0645   0.8527   0.0406
  -0.750   0.2422   0.01136   0.00322  -0.0644   0.8228   0.0404
  -0.500   0.2695   0.01115   0.00284  -0.0640   0.7933   0.0404
  -0.250   0.2961   0.01104   0.00252  -0.0635   0.7657   0.0406
   0.000   0.3228   0.01099   0.00228  -0.0631   0.7414   0.0412
   0.250   0.3495   0.01098   0.00209  -0.0627   0.7190   0.0423
   0.500   0.3761   0.01098   0.00194  -0.0623   0.6966   0.0449
   1.000   0.4258   0.00867   0.00191  -0.0610   0.6547   1.0000
   1.250   0.4522   0.00884   0.00191  -0.0606   0.6333   1.0000
   1.500   0.4786   0.00902   0.00193  -0.0603   0.6127   1.0000
   1.750   0.5049   0.00919   0.00196  -0.0600   0.5904   1.0000
   2.000   0.5312   0.00938   0.00202  -0.0597   0.5695   1.0000
   2.250   0.5578   0.00955   0.00212  -0.0594   0.5495   1.0000
   2.500   0.5843   0.00973   0.00222  -0.0592   0.5303   1.0000
   2.750   0.6109   0.00991   0.00234  -0.0590   0.5135   1.0000
   3.000   0.6376   0.01009   0.00248  -0.0588   0.4963   1.0000
   3.250   0.6642   0.01027   0.00267  -0.0586   0.4767   1.0000
   3.500   0.6904   0.01049   0.00285  -0.0583   0.4541   1.0000
   3.750   0.7166   0.01072   0.00305  -0.0581   0.4259   1.0000
   4.000   0.7404   0.01122   0.00326  -0.0575   0.3552   1.0000
   4.250   0.7613   0.01220   0.00371  -0.0568   0.2538   1.0000
   4.500   0.7849   0.01288   0.00416  -0.0564   0.2089   1.0000
   4.750   0.8094   0.01343   0.00456  -0.0560   0.1674   1.0000
   5.000   0.8314   0.01438   0.00510  -0.0555   0.1012   1.0000
   5.250   0.8514   0.01577   0.00613  -0.0544   0.0186   1.0000
   5.500   0.8757   0.01645   0.00698  -0.0538   0.0149   1.0000
   5.750   0.8994   0.01726   0.00800  -0.0530   0.0132   1.0000
   6.000   0.9214   0.01835   0.00935  -0.0520   0.0114   1.0000
   6.250   0.9442   0.01917   0.01034  -0.0512   0.0102   1.0000
   6.500   0.9649   0.02030   0.01164  -0.0501   0.0096   1.0000
   6.750   0.9841   0.02162   0.01313  -0.0488   0.0092   1.0000
   7.000   1.0026   0.02309   0.01471  -0.0474   0.0089   1.0000
   7.250   1.0213   0.02472   0.01644  -0.0459   0.0086   1.0000
   7.500   1.0402   0.02636   0.01815  -0.0447   0.0080   1.0000
   7.750   1.0572   0.02871   0.02050  -0.0436   0.0071   1.0000
   8.000   1.0759   0.03124   0.02321  -0.0423   0.0065   1.0000
   8.250   1.0949   0.03334   0.02558  -0.0410   0.0064   1.0000
   8.500   1.1120   0.03575   0.02830  -0.0396   0.0064   1.0000
   8.750   1.1265   0.03845   0.03132  -0.0380   0.0063   1.0000
   9.000   1.1380   0.04142   0.03470  -0.0361   0.0064   1.0000
   9.250   1.1459   0.04466   0.03829  -0.0342   0.0064   1.0000
   9.500   1.1500   0.04821   0.04217  -0.0321   0.0065   1.0000
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